Gas turbine engine control for rotor bore heating

ABSTRACT

A gas turbine engine comprises a compressor rotor including blades and a disk, with a bore defined radially inwardly of the disk, and a combustor. A tap directs the products of combustion first to a valve and then into the bore of the disk. At least two temperature sensors sense a temperature of the products of combustion downstream of the valve. A control compares sensed temperatures from the at least two temperature sensors to ensure the at least two temperature sensors are functioning properly. The sensed temperatures are utilized to control the valve. A method of operating a gas turbine engine is also disclosed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation-in-part of PCT Application NumberPCT/US2014/015759, filed Feb. 11, 2014, which claims the benefit of U.S.Provisional Application No. 61/766,321, filed Feb. 19, 2013.

BACKGROUND OF THE INVENTION

This application relates to a method of providing heat to a rotor borein a gas turbine engine at certain times during operation of anaircraft.

Gas turbine engines are known and when used on aircraft typicallyinclude a fan delivering air into a bypass duct and into a compressorsection. Air from the compressor is passed downstream into a combustionsection where it is mixed with fuel and ignited. Products of thiscombustion pass downstream over turbine rotors driving them to rotate.

Turbine rotors drive compressor and fan rotors. Historically, the fanrotor was driven at the same speed as a turbine rotor. More recently, ithas been proposed to include a gear reduction between the fan rotor anda fan drive turbine. With this change, the diameter of the fan hasincreased dramatically and a bypass ratio or volume of air deliveredinto the bypass duct compared to a volume delivered into the compressorhas increased. With this increase in bypass ratio, it becomes moreimportant to efficiently utilize the air that is delivered into thecompressor.

One factor that increases the efficiency of the use of this air is tohave a higher pressure at the exit of a high pressure compressor. Thishigh pressure results in a high temperature increase. The temperature atthe exit of the high pressure compressor is known as T₃ in the art.

There is a stress challenge to an increasing T₃ on a steady state basisdue largely to material property limits called “allowable stress” at agiven maximum T₃ level. At the maximum, a further increase in a designT₃ presents challenges to achieve a goal disk life. In particular, asthe design T₃ is elevated, a transient stress in the disk increasesbecause the radially outer portions of a high pressure compressor rotor(i.e., the blades and outermost surfaces of the disk or blisk), whichare in the path of air, see the increased heat rapidly when T₃ shoots uprapidly during a rapid power increase such as when the pilot increasespower during a take-off roll. However, a rotor disk bore does not seethe increased heat as immediately. Thus, there are severe stresses dueto the thermal gradient between the disk bore and the outer rim region.

This thermal gradient challenge is greatest during the take-off of anaircraft engine and it is possible that the thermal stress in the diskis much greater than the stress due to the centrifugal force on thedisk—particularly in the compressor where the blades are light. Theengine has typically been at low speed or idle as the aircraft waits onthe ground and then, just before take-off, the speed of the engine isincreased dramatically. The thermal gradient stresses have led to thehigh pressure compressor often being operated at a lower pressure (andhence T₃) than would be optimum.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine comprises a compressorrotor including blades and a disk, with a bore defined radially inwardlyof the disk, and a combustor. A tap directs the products of combustionfirst to a valve and then into the bore of the disk. At least twotemperature sensors sense a temperature of the products of combustiondownstream of the valve. A control compares sensed temperatures from theat least two temperature sensors to ensure the at least two temperaturesensors are functioning properly. The sensed temperatures are utilizedto control the valve.

In another embodiment according to the previous embodiment, the controlis programmed to compare the sensed temperatures as a check on theaccuracy of the at least two temperature sensors, and if the sensedtemperatures are not equal, then comparing the sensed temperatures to anestimated temperature based upon another engine parameter.

In another embodiment according to any of the previous embodiments, theanother engine parameter is rotor speed.

In another embodiment according to any of the previous embodiments, thevalve is closed if the estimated temperature is not equal to any of thesensed temperatures, and if one of the sensed temperatures is close tothe estimated temperature, then the temperature sensor associated withthe one of the sensed temperatures is utilized to control the valve.

In another embodiment according to any of the previous embodiments, thecontrol further senses a closed temperature of the supply line when thevalve is closed, and if the temperature exceeds a pre-determinedmaximum, then reducing power supplied by the engine.

In another embodiment according to any of the previous embodiments, thevalve is controlled to only be opened under certain periods of operationof the gas turbine engine.

In another embodiment according to any of the previous embodiments, thecontrol is programmed to be open when an aircraft associated with thegas turbine engine is at idle or low speed, and to be closed at take-offand higher speed operation.

In another embodiment according to any of the previous embodiments, asecond valve is provided to provide a redundant shutoff valve.

In another embodiment according to any of the previous embodiments,there are two of the taps, with each of the two taps including at leasttwo temperature sensors and at least two of the valves.

In another embodiment according to any of the previous embodiments, theproducts of combustion are also delivered to a bore of a turbinesection.

In another embodiment according to any of the previous embodiments, thetap is located within the combustor.

In another embodiment according to any of the previous embodiments, thetwo temperature sensors are mounted on a supply line receiving the tap.

In another embodiment according to any of the previous embodiments, thetemperature sensors are mounted downstream of an outlet of a supplyline.

In a further embodiment, a method of operating a gas turbine enginecomprises the steps of directing the products of combustion first to avalve and then into a bore of a compressor rotor disk. The temperatureof the directed products of combustion downstream of the valve aresensed with at least two temperature sensors. The sensed temperaturesfrom the at least two temperature sensors are compared to ensure the atleast two temperature sensors are functioning properly. The sensedtemperatures are utilized to control the valve.

In another embodiment according to the previous embodiment, the twosensed temperatures are compared as a check on the accuracy of bothsensed temperatures, and if the sensed temperatures are not equal, thenthe sensed temperatures are compared to an estimated temperature basedupon another engine parameter.

In another embodiment according to any of the previous embodiments,rotor speed is used as the another engine parameter.

In another embodiment according to any of the previous embodiments, thevalve is closed if the estimated temperature is not equal to either ofthe sensed temperatures, and if one of the sensed temperatures is closeto the estimated temperature, then the one of the sensed temperatures isutilized to control the valve.

In another embodiment according to any of the previous embodiments, atemperature of the supply line is sensed when the valve is closed, andif the temperature exceeds a pre-determined maximum, then power suppliedby the engine is reduced.

In another embodiment according to any of the previous embodiments, thevalve is only opened under certain periods of operation of the gasturbine engine.

In another embodiment according to any of the previous embodiments, thevalve is open when an aircraft associated with the gas turbine engine isat idle or low speed, and closed at take-off and higher speed operation.

In another embodiment according to any of the previous embodiments, asecond valve is provided to provide a redundant shutoff valve.

In another embodiment according to any of the previous embodiments,there are two of the taps, with each of the two taps including at leasttwo temperature sensors and at least two of the valves.

In another embodiment according to any of the previous embodiments, thetwo temperature sensors are mounted on a supply line receiving the tap.

In another embodiment according to any of the previous embodiments, thetemperature sensors are mounted downstream of an outlet of a supplyline.

In another embodiment according to any of the previous embodiments, theproducts of combustion in the combustor are directed first to the valveand then into the bore of the compressor rotor disk is performed withinthe combustor.

In another featured, a gas turbine engine comprises a compressor rotorincluding blades and a disk, with a bore defined radially inwardly ofthe disk, and a combustor including a burner nozzle. A tap directs theproductions of combustion in the combustor through a valve, and into thebore of the disk.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a schematic of a gas turbine engine.

FIG. 2 shows a detail of a method for allowing an increase in acompressor exit temperature.

FIG. 3A shows a control embodiment.

FIG. 3B shows a second control embodiment

FIG. 4 shows another embodiment of the tapping system.

FIG. 5 is a control flow chart.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a turbofan gas turbine engine in the disclosednon-limiting embodiment, it should be understood that the conceptsdescribed herein are not limited to use with turbofans as the teachingsmay be applied to other types of turbine engines including three-spoolarchitectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the high pressureturbine 54 and the low pressure turbine 46. The mid-turbine frame 57further supports bearing systems 38 in the turbine section 28. The innershaft 40 and the outer shaft 50 are concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A which iscollinear with their longitudinal axes.

To facilitate discussion of the engine, the fan and gear architectureare often referred to as the engine propulsor. The compressor section,combustor and turbine section, on the other hand, are often referred toas the gas generator. However, other component groupings and descriptorsmay be utilized without limiting the nature or scope of the disclosedembodiments.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gearsystem or other gear system, with a gear reduction ratio of greater thanabout 2.3 and the low pressure turbine 46 has a pressure ratio that isgreater than about 5. In one disclosed embodiment, the engine 20 bypassratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout 5:1. Low pressure turbine 46 pressure ratio is pressure measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of the low pressure turbine 46 prior to an exhaust nozzle.The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.5:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present invention is applicable toother gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

FIG. 2 shows an engine 100 incorporating a compressor section 102 havingrotor blades 103 and a disk 104. A bore area sits radially inwardly ofthe disk 104. The thermal gradient between an inner surface of disk 104and the outer region near the blades 103 is the stress challengementioned above. An “allowable stress” factors material strength at agiven temperature to achieve a preferred flight-cycle fatigue life forthe disk. For a commercial engine, the flight-cycle life might be 15,000flights. If T₃, or temperature at the exit of the high pressurecompressor, is increased while a disk is at its flight-cycle life, forexample, to increase rated Sea Level Takeoff thrust to power a heavieraircraft, than the disk stress will rise above the allowable stress andthe life of the disk may decrease.

As known, burner nozzles 114 are associated with a combustor 112. Inaccordance with this disclosure, a tap 116 removes hot air from thecombustor 112 and delivers it through a temperature regulating andshutoff valve 120. This air will be delivered into a chamber 130 topreheat disk 104 as will be described below.

The tap 116 may be jacketed in a pipe 118, carrying air from a plenumchamber 119 outward of the combustor 112. The jacket insulates thetapped air and keeps it at a relatively high temperature. Also, the pipeor tap 116 (or other plumbing assembly) near the area of the combustor112 will see temperatures well above two thousand (2000) degrees F.during take-off, and even up to this temperature at idle. It maytherefore be necessary to cool tap 116 to reduce peak temperatures, andprevent failures that stem from oxidation corrosion.

The plenum chamber 119 typically would have air at an exit temperatureof a high pressure compressor, or T₃ which might be above four hundred(400) degrees F. at idle and above one thousand (1000) degrees F. attakeoff. These temperatures would generally be at the point X₁ as shownin FIG. 2 and all around the combustor 112. At this time, inside thecombustor 112, the air temperatures are well below two thousand (2000)degrees F. at idle and well above this temperature degrees F. attakeoff.

In general, it is desirable to increase the pressure of the air reachingthe combustor 112, though this typically results in a higher temperatureT₃ at point X₁. Thus, a higher pressure at point X₁ results in anincrease in the thermal gradient challenge mentioned above. The air inthe combustor 112 tapped at 116 is at a temperature known in the art asT₄, or is the temperature leaving the combustor 112 and approaching afirst vane 106 of a turbine section 107. This is point X₂ of FIG. 2.

As shown, a turbine blade 108 is associated with the turbine disk 110.Similar to the above described thermal gradient challenge, with regardto that in the compressor, there is a thermal gradient challenge betweenthe regions of the disk holding the blade 108 and the disk 110 of therotor in the turbine section 107.

The hot air tapped from the combustor 112 may be over 1400 degrees whenan engine is at idle or low speed. This temperature may be within anacceptable range to preheat disks 104 and 110. It might also be muchhotter than 1400 degrees F. depending on the ambient temperature and theother demands on the engine such as cabin cooling bleed air being used,generator loads, and etc. Thus, even at idle, an unduly high temperaturemight be reduced by mixing in cooler air.

At takeoff power, the air in combustor 112 can approach temperatures of2700° F. during high speed operation. It would be undesirable to haveair at such high temperatures contacting disks 104 and 110. To remedythis problem, a relatively higher portion of cooling air is mixed withthe tapped hot air, and the duration at which the engine is at takeoffpower is monitored. During this process, once the disk 104 is calculatedto be at a desired preheat temperature, which will occur relativelyquickly, the preheat system is shut off by the valve 120.

The bore disk 109 is never exposed to the cool idle temperature, exceptperhaps during taxi or after landing, when the preheat system is notneeded. Note that valve 120 may also turn on the preheat air tap 116 ifthe engine is at low power for a long period. As an example, an aircraftmay operate at relatively low power, due to an air traffic delay aftertakeoff, and then the pilot may re-accelerate to climb power. Thetemperature regulating and shut off valve 120 is thus only open duringcertain periods of operation of the engine 100.

Downstream of the temperature regulating and shut off valve 120 there isa second valve 122. The second valve 122 is provided for failsafeoperation. As mentioned, it would be undesirable for the air from tap116 to leave the combustor and contact the disks 104, 110 during manyperiods of operation. The second valve increases the likelihood that atleast one valve 120 or 122 will function to prevent this air frompassing both valves 120 and 122 under certain conditions, as describedbelow. If a valve fails to operate, the engine control sets amaintenance flag. If both fail to close, the engine control reducespower or shuts the engine down for safety reasons

A small, motor driven, supplemental compressor 124 may be included in aflow path 127, if necessary for driving the air flow. However, thecompressor 124 may not be necessary depending on the overall designpressures of the engine and the overall design pressures of the heatingsystem. The flow path 127 may pass through a chamber 126 and opening 125with labyrinth seals to access the rotating inner drum area, and thenpass both upstream and downstream direction in the drum.

Air is shown passing upstream into chamber 130. The air in chamber 130passes through a rotating passage or pipe 132 and is directed againstthe radially inner surfaces of the disk 104. The air also heats otherareas that contact the air flow first, and the air loses heat all alongits path. A reason why this system starts out with hot combustion air isthat heat transfer typically reduces the air temperature along the way.

In addition, air flows downstream into chamber 128 and also heats theturbine disk 110, especially in the bore area.

As shown, a labyrinth seal 141 may be positioned in a chamber to preventleakage of the hot air toward a diffuser case 142 inwardly of the vane106.

The air tapped at 116 may be above 1400° F. at low power or idleoperation, and upwards of 2700° F. at takeoff. The air in the chamber119 may be on the order of 400° during low speed operation. Thedisclosed system preheats the disk 104 prior to the engine 100 enteringtakeoff mode at which the blades 103 will experience a rapid increase intemperature. Preheating the disk to, e.g., 800 degrees, F, allows the T₃temperature to be very high. Notably, it may be desirable to preheat thedisk 104 to a temperature that is actually higher than the T₃temperature at low speed operation.

In a method of operating an engine 100, a control 200 for the engine 100may direct hot air from tap 116 into the bore 130 of the compressor 130and into the turbine section 107 bore by opening valves 120 and 122.This flow may occur before take-off of an aircraft associated with theengine 110 and during low speed operation such as taxi or idle.

The valves 120 and 122 are then closed prior to take-off. Thus, theextreme high temperatures generated in the combustor 112 at take-off arenot delivered into the bores.

While tap 116 is in the combustor 112, it is possible to tap the airfrom other locations downstream of the combustor 112, say in the turbinesection 107. Such air has already been heated in the combustor 112.

With this preheat system, the disk 104 of the compressor is preheatedmuch closer to the temperatures the blade 103 reaches at take-off. Thisallows an increased pressure at point X₁. With higher pressure comeshigher temperatures, but this system allows the compressor to experiencesuch higher temperatures. The resulting higher pressures reachingcombustor 112 dramatically increase combustion efficiency.

In addition, it is desirable to have periodic self-tests on theoperation of the valves 120 and 122 to ensure that they are bothavailable and to ensure the redundant and failsafe operation.Accordingly, the system may cycle full open and full closed, forexample, immediately after engine start and the valve positions reportedto the electronic engine control 200.

FIG. 3A shows an alternate control 202 which may be incorporated into asystem such as shown in FIG. 2. The supply line 127 of FIG. 2 isreplaced by a supply line 201, which may be incorporated into a systemotherwise structured and controlled as in FIG. 2. Thermocouples 204 and206 are placed on the supply line 201, downstream of the valves. Whilethermocouples are disclosed, other temperature sensors may be used. Theredundant thermocouples ensure that the failure of one thermocouple willnot lead to a lack of accurate information with regard to thetemperature of the gases in the supply line 201. As mentioned above, itis important that these temperatures not exceed a pre-determinedmaximum.

While the thermocouples 204 and 206 are on the supply line 201 in theFIG. 3A embodiment, FIG. 3B shows an alternative embodiment wherein thesupply line 201B has an outlet, and the thermocouples 204B and 206B aredownstream of the outlet, and mounted within the engine. This embodimentmay reduce the complexity of providing electric communication.

FIG. 4 shows another embodiment 210 of an engine again centered on lineA. Engine 210 has two circumferentially spaced redundant taps 212 and214. Each may be structured as tap 116 (with all associated downstreamstructure), generally disclosed in FIG. 2. Again, the provision ofredundant supplies will ensure that a failure on either of the valvingsystems will not result in a failure to provide the heating. Rather,should one fail, one of the two valving systems can be shut down, withanother left operational.

FIG. 5 shows a control flow chart. In a step 220, the control 202 askswhether the temperature T_(A) from the thermocouple 304 is equal to thetemperature T_(B) from the thermocouple 206. While the step 220indicates the question is whether they are equal, it may simply be acomparison to ensure they are within some small percentage (for example5%) of each other. For purposes of interpreting the claims of thisapplication, the word “equal” should be interpreted this way.

If they are not equal, then at step 222, the control 202 estimates atemperature based on the engine speed T_(N). The T_(N) is compared toboth temperatures T_(A) and T_(B). If T_(N) is equal to one of the twotemperatures, then the T_(A) or T_(B) which is close to the estimatedT_(N) is utilized. If T_(N) is not equal to T_(A) or T_(B), then thevalves are closed at step 223.

If a system with redundant taps is utilized, such as in engine 210, oneof taps 212/214 may be closed while the other continues to supply gas toheat the bore.

Returning to step 220, if T_(A) is generally equal to T_(B), then thetemperatures are compared to the T_(N) to confirm the system is properlyfunctioning at step 224.

At step 228, the system also asks whether when the valves are closed,with the T_(A) and/or T_(B) still being utilized, and whether the sensedtemperature is above some maximum temperature (and here 1000° F.). If itis, then the power for the system is reduced. Alternatively, the enginemay be shut down. However, the redundant controls may allow the engineto continue to operate to provide some thrust even though there is aconcern indicated at step 228 by lowering power.

As can be appreciated, the control ensures the temperature reaching thecompressor disk bore will not be unduly high. Several valves can becontrolled to mix the tapped air to achieve a desired temperature. Ofcourse, a control can also shut the valves down should the temperaturebe so high that it cannot be utilized.

Should any of the thermocouples be found to be faulty, a maintenanceflag is set. Further, the entire system should be periodicallymonitored, and a full system health check provided periodically, such asevery 10^(th) ground idle operation.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

1. A gas turbine engine comprising: a compressor rotor including bladesand a disk, with a bore defined radially inwardly of said disk; acombustor; a tap for directing the products of combustion first to avalve and then into said bore of said disk; and at least two temperaturesensors for sensing a temperature of said products of combustiondownstream of said valve, and a control comparing sensed temperaturesfrom said at least two temperature sensors to ensure said at least twotemperature sensors are functioning properly, said sensed temperaturesbeing utilized to control said valve.
 2. The gas turbine engine as setforth in claim 1, wherein the control is programmed to compare thesensed temperatures as a check on the accuracy of the at least twotemperature sensors, and if the sensed temperatures are not equal, thencomparing said sensed temperatures to an estimated temperature basedupon another engine parameter.
 3. The gas turbine engine as set forth inclaim 2, wherein said another engine parameter is rotor speed.
 4. Thegas turbine engine as set forth in claim 2, wherein said valve is closedif said estimated temperature is not equal to any of said sensedtemperatures, and if one of said sensed temperatures is close to saidestimated temperature, then the temperature sensor associated with saidone of said sensed temperatures is utilized to control said valve. 5.The gas turbine engine as set forth in claim 4, wherein said controlfurther sensing a closed temperature of said supply line when said valveis closed, and if said temperature exceeds a pre-determined maximum,then reducing a power supplied by said engine.
 6. The gas turbine engineas set forth in claim 1, wherein said valve is controlled to only beopened under certain periods of operation of the gas turbine engine. 7.The gas turbine engine as set forth in claim 6, wherein the control isprogrammed to be open when an aircraft associated with the gas turbineengine is at idle or low speed, and to be closed at take-off and higherspeed operation.
 8. The gas turbine engine as set forth in claim 1,wherein a second valve is provided to provide a redundant shutoff valve.9. The gas turbine engine as set forth in claim 8, wherein there are twoof said taps, with each of said two taps including at least two saidtemperature sensors and at least two of said valves.
 10. The gas turbineengine as set forth in claim 1, wherein the products of combustion arealso delivered to a bore of a turbine section.
 11. The gas turbineengine as set forth in claim 1, wherein the tap is located within thecombustor.
 12. The gas turbine engine as set forth in claim 1, whereinsaid two temperature sensors are mounted on a supply line receiving saidtap.
 13. The gas turbine engine as set forth in claim 1, wherein saidtemperature sensors are mounted downstream of an outlet of a supplyline.
 14. A method of operating a gas turbine engine comprising thesteps of: directing the products of combustion first to a valve and theninto a bore of a compressor rotor disk; and sensing the temperature ofsaid directed products of combustion downstream of said valve with atleast two temperature sensors, comparing the sensed temperatures fromsaid at least two temperature sensors to ensure said at least twotemperature sensors are functioning properly, said sensed temperaturesbeing utilized to control said valve.
 15. The method as set forth inclaim 14, further including the step of comparing the two sensedtemperatures as a check on the accuracy of both sensed temperatures, andif the sensed temperatures are not equal, then comparing said sensedtemperatures to an estimated temperature based upon another engineparameter.
 16. The method as set forth in claim 15, further includingthe step of using rotor speed as said another engine parameter.
 17. Themethod as set forth in claim 16, further including the step of closingsaid valve if said estimated temperature is not equal to either of saidsensed temperatures, and if one of said sensed temperatures is close tosaid estimated temperature, then said one of said sensed temperatures isutilized to control said valve.
 18. The method as set forth in claim 14,further including the step of sensing a temperature of said supply linewhen said valve is closed, and if said temperature exceeds apre-determined maximum, then reducing a power supplied by said engine.19. The method as set forth in claim 14, wherein said valve is onlyopened under certain periods of operation of the gas turbine engine. 20.The method as set forth in claim 19, wherein said valve is open when anaircraft associated with the gas turbine engine is at idle or low speed,and closed at take-off and higher speed operation.
 21. The method as setforth in claim 14, wherein a second valve is provided to provide aredundant shutoff valve.
 22. The method as set forth in claim 21,wherein there are two of said taps, with each of said two taps includingat least two said temperature sensors and at least two of said valves.23. The method as set forth in claim 14, wherein said two temperaturesensors are mounted on a supply line receiving said tap.
 24. The methodas set forth in claim 14, wherein said temperature sensors are mounteddownstream of an outlet of a supply line.
 25. The method set forth inclaim 11, wherein the step of directing the products of combustion inthe combustor first to the valve and then into the bore of thecompressor rotor disk is performed within the combustor.
 26. A gasturbine engine comprising: a compressor rotor including blades and adisk, with a bore defined radially inwardly of said disk; a combustor,including a burner nozzle; and a tap for directing the productions ofcombustion in the combustor through a valve, and into said bore of saiddisk.